1. Field of the Invention
The present invention relates to improvements in solid propellant rocket motors or gas generators, and more particularly, to such devices having a plurality of solid propellant units disposed therein and embodying a membrane seal structure that enables the ignition of the propellant units to be independent of each other whereby discrete impulses are available upon command.
2. Description of the Prior Art
The entire propulsive capability of solid propellant rocket motors is usually spent during the combustion process of one mass of solid propellant. This is for the reason that once a solid propellant is ignited it is very difficult to stop the combustion process until the entire mass of ignited propellant has been consumed.
It has been proposed in the prior art to provide a solid propellant rocket motor with an ability to fire more than once, that is, a rocket motor with a "start-stop-restart" capability by providing two or more concentric units, that is, layers of zones of solid propellant in a combustion chamber with a flame inhibiting barrier separating the layers, the barrier being made of a material that will confine the burning to a single layer or zone but nevertheless is destructible so that the next adjacent layer may be ignited.
One such prior art arrangement for providing a rocket motor that may be fired more than once is disclosed in U. S. Pat. No. 3,293,855 granted on Dec. 27, 1966 to W. E. Cutill et al wherein a pyrotechnic and an electrically ignitable film are provided between each of the layers for igniting, upon command, and in turn, each of the next adjacent layers.
Other such prior art arrangements are disclosed in U. S. Pat. Nos. 3,564,845 granted to I. H. Friedman et al, Jr. on Feb. 23, 1971 and 3,568,448 granted to G. E. Webb, Jr. on Mar. 9, 1971 wherein one of two solid propellant concentric layers that are separated by a flame inhibiting barrier is ignited by an igniter that is extended through the rocket motor nozzle into the combustion chamber. The other layer is ignited by a gas generator that is connected by a tubular extension to the head end of the combustion chamber. A rupturable membrane seal and perforated support member assembly is provided to isolate the gas generator from the motor combustion chamber during burning of the first rocket propellant layer.
U.S. Pat. Nos. 3,340,691 granted on Sept. 12, 1967 to G. F. Mangum and 3,354,647 granted on Nov. 28, 1967 to W. C. Aycock disclose similar arrangements but involve the admission of liquid fuel to the combustion chamber for the destruction of the flame inhibiting barrier and the ignition of the adjacent layer of propellant.
All of such prior art patented disclosures are characterized in the provision of a single combustion chamber for a plurality of concentric solid propellant layers or zones, in involving the destruction of the flame inhibiting barrier between layers in order to initiate combustion and firing of the next adjacent layer, and in the use of a relatively complicated ignition arrangement for effecting destruction of the barrier and ignition of the next adjacent layer. The Friedman, Jr. et al and Webb, Jr. patented disclosures further involve the use of a rupturable membrane and perforated support member which when ruptured and disintegrated, respectively, tend to introduce debris into the combustion chamber. The Mangum and Aycock patented disclosures further involve the admission of liquid fuel to the combustion chamber.